Adaptive Variable Structure Controller for Spacecraft Vibration Reduction

A hybrid control scheme to vibration reduction of flexible spacecraft is presented by using a variable structure technique for attitude control and piezoelectric materials for active vibration suppression. The attitude controller consists of a linear feedback term and a discontinuous feedback ones. An adaptive version of the proposed attitude controller is also achieved through releasing the limitation of knowing the bounds of the lumped perturbations in advance. An additional independent control system acting on the flexible parts can be designed for further vibration suppression. Simulation results have shown the theoretical and practical merit of this approach.     

A composite control scheme for attitude maneuvering and elastic mode stabilization of flexible spacecraft with measurable output feedback

This paper treats the question of attitude maneuver control and elastic mode stabilization of a flexible spacecraft based on adaptive sliding mode theory and active vibration control technique using piezoelectric materials. More precisely, a modified positive position feedback (PPF) scheme is developed to design the PPF compensator gains in a more systematical way to stabilize the vibration modes in the inner loop, in which a cost function is introduced to be minimized by the feedback gains subject to the stability criterion at the same time. Based on adaptive sliding mode control theory, a discontinuous attitude control law is derived to achieve the desired position of the spacecraft, taking explicitly into account the mismatched perturbation and actuator constraints. In the attitude control law, an adaptive mechanism is also embedded such that the unknown upper bound of perturbation is automatically adapted. Once the controlled attitude control system reaches the switching hyperplane, the state variables can be driven into a small bounded region. An additional attractive feature of the attitude control method is that the structure of the controller is independent of the elastic mode dynamics of the spacecraft, since in practice the measurement of flexible modes is not easy or feasible. The proposed control strategy has been implemented on a flexible spacecraft. Both analytical and numerical results are presented to show the theoretical and practical merit of this approach.     

应用分力合成主动振动抑制方法的最优飞行器大角度机动控制策略

 针对当代带大型挠性附件的空间飞行器,提出了分力合成主动振动抑制方法,并且分析了方法的鲁棒性。该方法可以保证挠性飞行器在实现指定的刚体运动的同时,抑制掉对系统影响较大的挠性振动模态,对频率不确定性的鲁棒性使得该方法易于工程实践。对于使用常幅值力矩喷气执行机构的航天器,设计了应用分力合成方法的时间—燃料最优机动控制律,数值仿真结果验证了方法的有效性。     

柔性航天器振动主动抑制及姿态控制

针对柔性航天器柔性附件振动主动抑制及姿态高精度快速稳定问题，研究了一种输入成形器（IS）-自适应有限时间干扰观测器（FDO）-有限时间积分滑模控制器综合的设计方法。首先，基于柔性模态的频率及阻尼信息，获得能够有效抑制柔性振动的输入成形器形式，并与系统参考输入进行卷积，得到期望参考输入；其次，基于航天器动力学模型，设计一种新型的自适应有限时间干扰观测器，避免了综合干扰上界必须已知的约束，且保证干扰估计误差有限时间收敛至零，实现对干扰及残余振动影响的快速精确估计；最后，基于观测器的估计值，设计多变量有限时间积分滑模控制器，保证对期望参考输入的高精度快速跟踪控制，并进行严格的稳定性证明。仿真结果表明，该综合设计策略能够保证柔性附件振动抑制75%，姿态稳定度达到10^-4数量级。     

挠性飞行器姿态机动脉冲调宽控制系统设计

大挠性充液飞行器喷气姿态控制系统设计是一个复杂和关键的问题,为了抑制挠性振动,设计形状输入信号,选取双曲正切作为平滑函数,并和Bang-Bang控制对比研究.喷气脉冲调制采用工程上易于实现的脉冲调宽（PWM）设计方法,结合积分死区原理设计脉冲宽度,对具有挠性充液动力学特性的对象进行控制,抑制了振动模态和晃动模态,实现了平滑机动控制.通过数学仿真表明,形状输入法和合理的脉冲调宽逻辑可以有效抑制大挠性飞行器姿态机动控制过程中的挠性振动.     

基于输出反馈的柔性航天器变结构跟踪控制方法

为解决柔性航天器姿态机动的控制问题,给出了基于输出反馈的变结构跟踪控制算法。针对柔性航天器的大角度机动,在建立了柔性航天器相对参考轨迹的动力学方程的基础上,设计了仅利用航天器本体的角度和角速度信息的变结构跟踪控制器,使得姿态状态跟踪误差(包括姿态跟踪误差和姿态角速度跟踪误差)以及挠性附件的模态变量从任意的初始状态出发都会到达包含原点的一个闭集内,并且姿态状态跟踪误差能收敛到零,并给出了严格的数学证明。仿真结果证明了所提控制方法的可行性和有效性。     

空间飞行器大角度机动飞行的变结构姿态控制

研究了空间飞行器大角度机动飞行的变结构姿态控制。应用四元数来描述姿态运动，以消除大角度机动飞行时欧拉角描述所存在的奇异性。基于Lyapunov方法设计了变结构控制的切换函数，以保证系统的滑动模态．亦即四元数偏差的稳定性。基于所给切换函数，设计了变结构控制器。数值仿真的结果说明了设计方法的有效性。     

Adaptive control and stabilization of elastic spacecraft

This work treats the question of large angle rotational maneuver and stabilization of an elastic spacecraft (spacecraft-beam-tip body configuration). It is assumed that the parameters of the system are completely unknown. An adaptive control law is derived for the rotational maneuver of the spacecraft. Using the adaptive controller, asymptotically decoupled control of the pitch angle of the space vehicle is accomplished, however this maneuver causes elastic deformation of the beam connecting the orbiter and tip body. For the stabilization of the zero dynamics (flexible dynamics), a stabilizer is designed using elastic mode velocity feedback. In the closed-loop system including the adaptive controller and the stabilizer, reference pitch angle trajectory tracking and vibration suppression are accomplished. Simulation results are presented to show the maneuver capability of the control system     

Sliding mode and shaped input vibration control of flexible systems

In this paper, the vibration reduction problem is investigated for a flexible spacecraft during attitude maneuvering. A new control strategy is proposed, which integrates both the command input shaping and the sliding mode output feedback control (SMOFC) techniques. Specifically, the input shaper is designed for the reference model and implemented outside of the feedback loop in order to achieve the exact elimination of the residual vibration by modifying the existing command. The feedback controller, on the other hand, is designed based on the SMOFC such that the closed-loop system behaves like the reference model with input shaper, where the residual vibrations are eliminated in the presence of parametric uncertainties and external disturbances. An attractive feature of this SMOFC algorithm is that the parametric uncertainties or external disturbances of the system do not need to satisfy the so-called matching conditions or invariance conditions provided that certain bounds are known. In addition, a smoothed hyperbolic tangent function is introduced to eliminate the chattering phenomenon. Compared with the conventional methods, the proposed scheme guarantees not only the stability of the closed-loop system, but also the good performance as well as the robustness. Simulation results for the spacecraft model show that the precise attitudes control and vibration suppression are successfully achieved.     

太阳帆航天器动力学建模与求解

 太阳帆航天器动力学建模与求解是姿态控制与结构振动抑制的基础,具有重要的理论与工程意义。针对带有控制杆和控制叶片的太阳帆航天器,进行结构的合理简化。应用矢量力学基本原理,推导出考虑弹性振动的太阳帆航天器姿态动力学方程,再对其进行简化,分别得到基于控制叶片和控制杆的两类太阳帆航天器的姿态动力学方程,联立太阳帆支撑杆振动方程,结合非约束模态的定义对运行于超地球同步转移轨道的太阳帆航天器动力学方程进行了求解及分析,结果表明所建立的太阳帆动力学模型可准确地描述柔性太阳帆航天器的动力学特性。     

挠性航天器的退步直接自适应姿态跟踪控制

针对参数不确定的挠性航天器姿态跟踪控制问题,提出了一种退步直接自适应控制算法。首先验证了挠性航天器动力学子系统的近似严格正实性,并设计了具有理想控制性能的参考模型;然后对以姿态四元数描述的运动学子系统设计常系数输出反馈中间控制律,使航天器姿态四元数输出渐近跟踪参考模型输出;最后退一步,对具有参数不确定特性的动力学子系统,基于非线性直接自适应控制理论和Lyapunov稳定性理论,设计了退步直接自适应姿态跟踪控制器,并证明了闭环系统的稳定性。仿真结果表明,所提控制方法能有效抑制挠性附件的振动,对挠性航天器的控制是有效的。     

Acceleration sensors based modal identification and active vibration control of flexible smart cantilever plate

Some flexible appendages of spacecrafts, such as solar panels, are cantilever plate structures. Thus, vibration problem is unavoidable when there is slewing maneuver or external disturbance excitation. Vibration of such cantilever plate structures includes coupled bending and torsional motion. Furthermore, the low amplitude vibration near the equilibrium point is very difficult to be quickly suppressed due to nonlinear factors of the hardware in the system, which is harmful to stability and attitude control accuracy. To solve these problems, acceleration sensor-based modal identification and active vibration control methods are presented for the first two bending and the first two torsional modes vibration of the cantilever plate. Optimal placements of three acceleration sensors and PZT patches actuators are performed to decouple the bending and torsional vibration of such cantilever plate for sensing and actuating, and identifications are achieved by experiments. A nonlinear control method is presented to suppress both high and low amplitude vibrations of flexible smart cantilever plate significantly. Experimental comparison researches are conducted by using acceleration proportional feedback and the presented nonlinear control algorithms. The experimental results demonstrate that the presented acceleration sensor-based methods can suppress the vibration of cantilever plate effectively.     

敏捷卫星的联合执行机构控制策略

 针对对地观测敏捷卫星大角度快速机动、高控制精度的任务需求,提出了联合推力器与飞轮作为执行机构的控制策略。该控制策略综合利用2种执行机构的优点:推力器以前馈的形式提供机动过程中所需的主要力矩以实现航天器大角度的快速机动,而飞轮以反馈的形式提供精准的控制力矩以提高机动过程中的姿态控制精度。为补偿由于初始状态偏差和推力器输出力矩不准确所带来的控制误差,采用变结构控制设计了2种姿态跟踪控制器,使航天器能够渐进地跟踪上参考轨迹。并对姿态机动控制过程中,飞轮力矩及转速可能出现的饱和问题作了相应的修正。仿真结果表明了所提控制策略及所设计控制算法的可行性和有效性。     

Fault-tolerant control and vibration suppression of flexible spacecraft: An interconnected system approach

This paper considers a fault-tolerant control and vibration suppression problem of flexible spacecraft. The attitude dynamics is modeled by an interconnected system, in which the rigid part and the flexible part are coupled with each other. Such a model allows us to use the interconnected system approach to analyze the flexible spacecraft. Both distributed and decentralized observer-based fault-tolerant control schemes are developed, under which the closed-loop stability of flexible spacecraft can be ensured by using the cycle-small-gain theorem. Compared with the traditional method, this paper considers the faults occurred not only in the rigid parts, but also in the flexible parts. In addition, the application of the interconnected system approach simplifies the system model of flexible spacecraft, thereby the difficulty of theoretical analysis and engineering practice of fault-tolerant control of flexible spacecraft are greatly reduced. Simulation results show the effectiveness of the proposed methods and the comparison of different fault-tolerant control approach.     

Rotational maneuver of nonlinear uncertain elastic spacecraft

The question of attitude control and elastic mode stabilization of a spacecraft (orbiter) with beam-tip-mass-type payloads is considered. A three-axis moment control law is derived to control the attitude of the spacecraft. The derivation of the control moments acting on the spacecraft does not require any information on the system dynamics. The control law includes a reference model and a dynamic compensator in the feedback path. For damping out the elastic motion excited by the slewing maneuver, an elastic mode stabilizer is designed. The stabilization is achieved by modal velocity feedback using force and torque actuators located at the payload end of the elastic beam. Collocated actuators and sensors provide robust stabilization. Simulation results are presented to show that rotational maneuvers and vibration stabilization can be accomplished in the closed-loop systems despite the presence of model uncertainty and disturbance torque in the system     

基于干扰观测器的天基观测航天器姿态控制设计

与地基空间目标监视系统相比,天基观测系统具有监视范围广,不受国界限制,观测精度高等优点,是未来空间目标观测技术的重要发展方向。但天基观测航天器工作时,相机转台的运动,太阳能帆板挠性部件的弹性振动与航天器的姿态运动相互影响,构成强耦合的非线性系统,传统的控制方案无法实现对这类天基观测航天器的高精度姿态控制。文章针对某一空间观测航天器的任务要求,设计了基于干扰观测器的前馈补偿航天器姿态控制系统,仿真实验结果表明:姿态角控制精度小于 0.06°,姿态角速度精度小于 0.03(°)/s,达到了精度要求。     

推力器故障的刚体航天器自适应变结构容错控制

 针对刚体航天器存在未知惯量参数、推力器故障以及控制受限的姿态控制问题,提出了一类自适应变结构容错控制方法,显式地引入推力器输出的饱和幅值,以确保控制输出在其要求界的范围内;同时,引入控制参数在线自适应调整技术,提高了控制律对参数、干扰以及故障变化的自适应能力;对设计者而言,推力器故障信息不需要进行在线检测和分离。此外,进一步考虑存在推力偏差对系统性能的影响,设计控制器参数使得闭环系统对这类推力偏差具有L₂增益稳定性。最后,将设计的控制器应用于航天器的姿态机动控制,仿真结果表明该控制器能有效地抑制外部干扰、参数不确定性和推力器各种故障的约束,在完成姿态机动的同时,保证其控制输出满足饱和受限界的要求。     

三轴轮控航天器姿态大角度机动控制器设计

研究了飞轮作为执行机构的航天器姿态大角度机动问题。建立了飞轮控制的姿态动力学模型;基于Lya-punov稳定性定理,设计了闭环控制器并给出了参数自适应律的解析表达式;分析了有界噪声作用下该模型的控制效果。仿真结果表明,所设计的控制器具有较强的鲁棒性和抗干扰能力。     

基于新型终端滑模的航天器执行器故障容错姿态控制

针对受干扰的刚体航天器冗余执行器存在故障与控制受限的姿态跟踪控制问题,提出一类基于新型指数形式的非奇异快速滑模面（ENFTSM）与趋近律的姿态容错控制器设计方法。当部分推力器发生故障时,假设剩余推力器具有输出饱和特性且能提供足够推力保证航天器执行任务,相比一般终端滑模控制器,本文设计的控制器不仅能使系统状态以更快的速度到达平衡点,且不需要在线对执行器故障信息进行检测和分离。基于Lyapunov方法证明本文设计的控制器能保证闭环系统稳定,且能有效地抑制外部干扰、控制受限和执行器故障等约束。最后对提出的控制算法进行了数值仿真,其结果表明了该控制器的有效性。     

基于新型终端滑模的航天器执行器故障容错姿态控制简

 针对受干扰的刚体航天器冗余执行器存在故障与控制受限的姿态跟踪控制问题,提出一类基于新型指数形式的非奇异快速滑模面（ENFTSM）与趋近律的姿态容错控制器设计方法。当部分推力器发生故障时,假设剩余推力器具有输出饱和特性且能提供足够推力保证航天器执行任务,相比一般终端滑模控制器,本文设计的控制器不仅能使系统状态以更快的速度到达平衡点,且不需要在线对执行器故障信息进行检测和分离。基于Lyapunov方法证明本文设计的控制器能保证闭环系统稳定,且能有效地抑制外部干扰、控制受限和执行器故障等约束。最后对提出的控制算法进行了数值仿真,其结果表明了该控制器的有效性。